For the attitude stabilization of spacecraft, it is known to correct deviations from the desired attitude by means of jet impulses as soon as a predetermined permissible value of the control error is exceeded; in this regard, see U.S. Pat. No. 3,984,071.
In spacecraft which move in orbits around the earth, deviations from the desired attitude are usually measured by means of infrared horizon sensors (IRS), radio frequency sensors or sun sensors. In special cases, such as, for example, in scientific or earth reconnaissance satellites, gyroscopes and star sensors are also used.
Depending on the type of application, the reaction engines which are used can either be cold gas engines, for example, nitrogen from pressure reservoirs, or chemical engines, for example, hydrazine with catalytic decomposition, or bipropellant systems with fuel and oxidator.
For converting the substantially continuous deviation into appropriate pulse-shaped signals for the moments about the axes of the vehicle, conventionally so-called "pseudorate" or "pulse width pulse frequency" (PWPF) modulators are used. The corresponding control networks consist of lag filters for suppressing the sensor noise in the case of the pseudorate modulator or of proportional-differential (PD) units with filters in the case of the PWPF modulator, see, for example, J. E. Vaeth: "Compatibility of Impulse Modulation Techniques with Attitude Sensor Noise and Spacecraft Maneuvering", IEEE Transactions on Automatic Control, January 1965, pages 67 ff.
Both control configurations are well suited for the stabilization of vehicles, particularly satellites, whose construction does not permit any significant structural-elastic deformations, such as bending vibrations and torsional vibrations.
With the increasing spatial extension of such vehicles, their high degrees of slenderness in certain axial directions and their lightweight construction, such structural vibrations, which, naturally, also have very weak internal damping, are unavoidable and are of decisive importance for the dynamics.
For example, in three-axis stabilized communications satellites, the electrical energy for the devices used aboard and the payload for communications is usually generated by two-oppositely arranged solar generators which can be unfolded and conventionally have a width of 1 to 2 m, but can have a length of 5 to 20 and more meters in the unfolded state, while having a thickness of only a few centimeters. The frequencies of the lowermost natural bending and torsional natural vibrations of such structures are in the order of magnitude of 0.1 to 2 Hz, i.e., in a frequency range in which during certain disturbance torque conditions the above-mentioned modulators are also operating which causes the excitation of structural vibrations. The frequencies of the pure rigid body motions are generally about one order of magnitude lower than those of the structural vibrations.
By means of the controller concept of the conventional type, it is very difficult and, with increasing size of the spacecraft, even impossible to meet the partially contradicting requirements for accuracy of pointing satisfactory rigid body dynamics and simultaneous active structural vibration damping. The adjustable parameters of the regular system are partially fixed from the outset or severely limited in their selection by the requirements or physical properties of the components; for example:
the controller amplification by the required attitude accuracy at a given maximum disturbance torque;
the magnitude of the minimum jet pulse (thrust.times.nozzle opening time) by the delay of the valves and/or the physics of the combustion process;
the maximum number of jet activations over the service life (of, e.g., 7 to 10 years) by reliability requirements (wear, aging);
conditions of the controller time lead-lag ratio filter time constant) by the signal/noise ratio of the sensors.
In the past, it was still possible in part, in spacecraft having a predominantly rigid structure or only one dominant form of bending vibrations, to find an acceptable compromise in the selection of the adjustment parameters of the control system of the above-mentioned configurations: see, for example, AIAA Paper No. 76-266 "Attitude Stability of Flexible Spacecraft Which Use Dual Time Constant Feedback Network Pseudorate Control", 6th Communications Satellite Systems Conference, Montreal, Apr. 5-8, 1976. This has become impossible to an increasing extent because of the accuracy requirements which require the controller amplification to be further increased. Also, due to the structures becoming larger and, therefore, "softer", several forms of bending vibrations of various natural frequency for each vehicle axis must be taken into consideration and stabilized.
Under these conditions, employing stabilization concepts of conventional configuration, even with the best possible selection of the adjustable parameters, instabilities or at least non-linear vibrations, so-called limit cycles, are created which lead to an actuation of the reaction jets in the rhythm of the natural frequencies of structural vibrations and result in an unduly high fuel consumption and in structural stresses which may endanger the mission.
The objective of the present invention consists in specifying improved arrangements for the attitude stabilization of flexible vehicles, such as aircraft and spacecraft, which exhibit one or several weakly damped structural vibrations at different harmonic frequencies, and in which discontinuously operating actuators are used, like for instance reaction jets or gimballed momentum wheels with steplike positioning of the gimbals to produce control torques.
In other words the improvement aims at a considerable increase of the attitude accuracy simultaneously mastering the structural oscillation problems. To achieve this goal novel control concepts are required, which provide a sufficiently large number of degrees of freedom for control, which are largely independent from each other and can be freely selected in order to meet the various control requirements simultaneously.
This task is solved by an observer of at most third order per vehicle axis, a state regulator, and a modulator loop with relay characteristics, where, according to one preferred embodiment of the invention, the modulator loop does not imply a separate modulator network and internal feedback, but the feedback path from the modulator output into the observer additionally provides the function of these two modulator elements.
Advantage of the invention can be taken in all control systems where the characteristic features of resonance properties of the control plant at particular natural frequencies and discontinuously operating actuators apply. This is the case for instance in slender, i.e. flexible rockets and missiles moving in the earth athmosphere, implying discontinuous rudder control and in launch vehicles, particularly upper stages of these, as well as in magnetically levitated and guided vehicles of elastic construction or respectively with elastic suspension of levitation and guidance magnets, which move along elastically deformable rails, and in which discontinuously operating magnet current drivers are used.
Instead of reaction jets in spacecraft and satellites, it is also possible to use, for example, momentum wheels with single or double cardanic mounting and with step positioning of the frame or housings, or so-called "control moment gyros" and the like.